1. Field of the Invention
The present invention relates to automatic flight control systems for aircraft and, more particularly, to monitored fail passive dual channel aircraft flight control systems for positioning the attitude control surfaces thereof.
2. Description of the Prior Art
As is known, aviation regulatory agencies impose safety requirements on the performance of aircraft automatic flight control systems. For example, with regard to the aircraft pitch axis, should a malfunction occur, the change in normaL acceleration of the aircraft must be limited to values which are typically in the range of plus or minus one g. Conventionally, accurate servomotor torque limiting is utilized to limit the servomotor torque to a value such that the maximum permitted normal acceleration or load factor will not be exceeded at the most critical flight condition of the aircraft. Typically, servo authority limit is established on the basis of the high speed regime of the aircraft with an aft center of gravity. Present day requirements in some instances are even more stringent and require that the maximum load factor not be exceeded due to a hardover malfunction in a particular direction when the automatic pilot is holding the pitch attitude control surface at its maximum limit in the opposite direction, as might occur due to a mistrimmed craft condition. As conventional torque limiting is designed for a single flight condition, the prior art practice severely limits performance at other flight conditions. In general, performance under such conditions which require greater servo authority than that provided by the torque limit is seriously compromised. Typically, such a compromise condition would exist at low speeds with a forward center of gravity such as under landing approach conditions. Thus, although excessive load factor maneuvers are prevented, performance of the aircraft over portions of its flight envelope may be seriously jeopardized.
Performance of such prior art systems are therefore sensitive to changes in aircraft configuration such as changes in speed, center of gravity position, flap-slat position, horizontal stabilizer position, changes in variable geometry aerodynamic control surface configurations and the like. Some prior art systems utilize monitor circuits to enable increased torque to be utilized; such systems are, however, sensitive to nuisance disconnects or false alerts.
Dual channel servo actuator systems are known in the prior art. The present invention, moreover, is an improvement of the monitored dual channel servo system disclosed in U.S. Pat. No. 4,035,705 to H. Miller entitled "Fail-Safe Dual Channel Automatic Pilot with Maneuver Limiting" which issued on July 12, 1977, to the present assignee of the instant disclosure. This patent discloses two servo channels, the mechanical outputs of which are coupled through a differential mechanism to position a control surface in a manner similar to the present disclosure. The apparatus of this patent reacts to the aircraft outer loop command source by processing the separate command signals through redundant displacement and rate limiting devices. In the pitch axis, separate normal accelerometers control each channel's command rate limit so as to prevent excessive pitch maneuvers due to the command source. Redundant polarity comparators are included for comparing the polarity or the incremental load factor with the polarity of each servo channel output and for applying a brake to clamp the output of a failed servo channel when the polarity of the incremental load factor is of the same sense as, i.e., agrees with the servomotor direction and the incremental load factor exceeds a predetermined level.
A further example of the prior art is U.S. Pat. No. 3,462,662 entitled "Monitoring and Fault Correction System for a Multiple Channel Servo Actuator" which issued on Aug. 19, 1969 in the name of W. E. Carpenter. Carpenter U.S. Pat. No. 3,462,662 is directed to a plurality of servo channels coupled through a differential mechanism in a manner similar to the present disclosure for actuation of a control surface. The monitor system of the Carpenter patent, however, reacts to a failure after one channel has failed to a saturated amplifier. That is, a difference threshold detector, in response to signals from the servo velocity generators thereof, which exceed a predetermined value that corresponds to the upper signal limit expected during normal operation, closes a set of relay contacts which couple a test signal from a test signal generator to the multiple servo channels. Inasmuch as one servo amplifier is in a saturated condition, the test signal applied to its respective channel does not alter the output condition of the motor associated therewith. However, the test signal applied to the operative channel is passed to the servo amplifier which acts upon it to either increase or decrease the output speed of the servomotor. Rate circuits detect the rate of change in the output signal from the velocity generators due to the test signals and, a threshold detector set at a predetermined value coupled to each rate circuit assures that the output thereof is not a spurious transient signal. The threshold detectors are then coupled through logic gate means to enable braking means to clamp the failed channel. Carpenter's U.S. Pat. No. 3,462,662 monitoring concept depends upon the reaction of the dual channels to a test signal applied thereto if a speed difference between the servoes is detected. Thus, in Carpenter, the braking of the failed channel can occur only after the failure is detected, a test signal is applied, the failed channel is identified and a threshold is exceeded.
Accordingly, it is necessary to provide a control system with means to rapidly detect a failure during the acceleration of the respective servomotor associated therewith such that any single failure will not cause significant movement or acceleration of the aircraft at any flight condition.